Gas turbine engines often include an annularly shaped combustor disposed radially outwardly of the primary rotating components of the engine. As such, the outer diameter of the annular combustor is often a critical determinant of the overall diameter of the gas turbine engine. Smaller diameter engines are not only critical in certain applications, but in addition can impact overall aircraft operations by reduction in drag.
FIG. 1 illustrates a typical annular combustor of the prior art and has an annular combustor case with annular inner and outer casings 14a, 16a. Within the plenum or combustor case is a combustor liner comprising inner and outer, annularly shaped, perforated liners 20a,22a. Pressurized air from the compressor suction of the engine is delivered through a diffuser 50a, and this pressurized air flow is split into a variety of flow paths into the combustion chamber 19a defined inside the combustor liner. A first portion of this airflow, approximately 23%, is delivered axially into the combustion chamber through a conventional airblast fuel nozzle 30a wherein fuel is finely atomized and initially mixed with the pressurized airflow. Outer annulus 15a carries approximately 20% of the airflow as primary airflow to support the combustion process, along with additional airflow that is introduced into the combustion chamber further downstream for purposes of cooling and dilution, as depicted by arrows 48a. Similarly, a portion of the pressurized airflow passes through the inner annulus 17a for delivery of primary airflow, comprising approximately 10% of the total air flow, along with additional downstream flow for cooling and dilution.
To carry the necessary airflow through the inner and outer annulus 17a,15a without introducing unacceptable back pressure on the airflow, the overall combustor is sized with a outer radius R.sub.a from the engine centerline 11a. This overall diameter R.sub.a is thus directly impacted by the necessary sizing of the inner and outer annulus spaces 17a and 15a.
In many applications the quantity of axial airflow passing through the transverse endwall 28a at the dome end of the combustor liner is limited because of flame stabilization problems, inadequate mixing for efficient combustion processes, and excessive thermal variations in the gas flow being exhausted from the combustion chamber.